Turbo fan jet propulsion jet engines for aircraft typically include compressor blades mounted on the forward end of engine rotors that rotate within a close fitting casing. The casing usually is secured to an overhead portion of the aircraft by a front mount assembly that depends downwardly from the aircraft to interconnect with the upper portion of the casing. Also, typically, a rear mount is employed that depends downwardly from the aircraft to interconnect a rearward section of the engine core. Spoke type struts extend radially outwardly from the engine forward core case to interconnect with the inside diameter of the fan casing.
The advent of large, high bypass turbo-fan jet propulsion engines has resulted in relative deflection between the engine casings and the rotors of engines that have been mounted to the aircraft by conventional means, thereby resulting in rubbing contact between the rotor blade tips and the engine casings. When the engines are operated at full power, such as during takeoff, the high thrust loads that act through the engine must be reacted by the engine thrust mounts. Since the engine thrust mounts are generally offset from the longitudinal center of the engine, bending moments are generated in the engine cases. The large bending loads resulting therefrom cause deflection of the engine components resulting in interference between the rotor blades and their associated casings.
In addition, when the aircraft is disposed at a high angle of attack, such as during takeoff, the air flowing around and through the inlet cowl pushes upwardly resulting in a bending moment being transferred from the cowling to the rearwardly interconnected fan casing and inwardly through the fan struts into the core casing. This bending moment results in further deflection of the casings and adds to the interference between rotating and stationary portions of the engine.
Excessive blade and tip seal wear increases the clearance between these components causing loss of fuel efficiency. In addition, a phenomenon known as blade tip stall may result from larger clearances between the blade tips and the engine casings. This can lead to vibrational problems associated with non-synchronous whirl motion of the rotors. Moreover, when compressor blades rub against their surrounding seal, particles are removed from the blade tips and the seal. The deposition of these particles on the extremely hot turbine sections of the engine roughens the turbine blades and stators and reduces their aerodynamic efficiency.
One standard technique for eliminating tip rub has been simply to provide larger clearances between the blades and the casing or selectively pregrind the sections of the casing which are susceptible to tip rub. Although this may assist in avoiding blade and seal particles from being deposited on the hot engine sections and avoiding loss of blade tip material, it results in lower fuel efficiency and can cause detrimental vibration of the rotors.
Another straighforward approach to reducing tip rubs is to thicken the engine casings and increase the number of rotor bearings to provide stiffening and better concentricity between rotors and stationary parts (casings, seals, frames). This approach, however, is very costly in terms of weight and complexity and is impractical in light of other alternatives.
Another attempted solution to the foregoing problems has been to mount the jet engine on an aircraft to avoid transfer of inlet induced loads through the engine core assembly. In this type of arrangement, the engine core assembly and the fan casing are mounted to a structural member of the aircraft, such as a pylon, independently of each other. However, since the pylon cannot be constructed in an absolutely rigid manner, under full engine power and at high angles of attack, a significant degree of relative movement may occurr between the engine core assembly and the fan casing leading to the rubbing of the fan blades against the casing. Examples of this type of engine mounting arrangement are disclosed by U.S. Pat. Nos. 3,750,983 and 4,013,245.
In a jet engine mounted in a nacelle structure, a proposed solution to the foregoing problems has been to place a hydraulic actuator between the bottom of the fan casing and the nacelle structure in an attempt to restrain the fan casing against movement relative to the nacelle structure. An example of this type of mounting structure is disclosed by U.S. Pat. No. 4,022,018.
Accordingly, it is a principal object of the present invention to provide a system for mounting a high bypass turbo fan jet propulsion engine to an aircraft so as to reduce blade and tip seal wear thereby increasing performance during takeoff and cruise, reducing fuel consumption and reducing deposition of metal particles from the compressor blades and tip seals on the very high temperature components of the engine.
More particularly, it is an object of the present invention to mount the high bypass turbo fan jet propulsion engine on an aircraft in such a manner to minimize the bending flexure and distortion of engine components caused by thrust loads and air loads on the engine inlet cowling.
Another object of the present invention is to achieve the foregoing objects while at the same time accommodating the thermal expansion and contraction of the engine components due to changes in the temperature of the engine.